Evolution and logistics of an early lunar base

The planned construction of a permanently manned space station in low earth orbit has reopened the discussion about the establishment of a manned lunar base within the next 25 years for exploration of the Moon and space. Several studies demonstrate that a lunar base very modest in size may cost $50 to 90 billion spread over 25 years which would fit into the expected NASA budget for this period. Having these cost in mind the authors present a concept having a greater effectiveness based on the following operational characteristics: (1) The development of a low cost heavy-lift launch vehicle for cargo transportation and propellant supply reduces the specific transportation cost by one order of magnitude compared to the existing Space Shuttle system. (2) Orbital transfer vehicles with LOX/LH2 technology should be preferred over advanced propulsion systems because of proved technology and cost reduction by utilization of lunar produced LOX. (3) The evolution of the lunar base towards a lunar colony and manufacturing facility could only be initiated by a powerful transportation system allowing for cost-effective space construction projects and manned spaceflight to other planets.The lunar base program of this paper is based on a schedule considering a 8 years development, 5 years lunar base assembly and 20 years operational phase during which the lunar crew will increase from 60 to 180 people. Launch rates will be 10 shuttle launches and 10 HLLV launches p.a. at the average. Development costs of the transportation and lunar base system will amount to $29 billion. Adding hardware and operational costs for lunar base assembly results in the acquisition cost of $49 billion. Total life cycle costs are estimated to be in the order of $101 billion considering a 20 years operational phase which will cost $2.6 billion p.a. at the average. For the 2508 man-years spent in lunosphere the relative cost will be $40.2 million per man-year of which space transportation will cost $25.0 million per man-year.     

An economically viable space power relay system

This paper describes and analyzes the economics of a power relay system that takes advantage of recent technological advances to implement a system that is economically viable. A series of power relay systems are described and analyzed which transport power ranging from 1,250 megawatts to 5,000 megawatts, and distribute it to receiving sites at transcontinental distances.

Two classes of systems are discussed—those with a single reflector and delivering all the power to a single rectenna, and a second type which has multiple reflectors and distributes it to 10 rectenna sites, sharing power among them. It is shown that when offering electricity at prices competitive to those prevalent in developed cities in the US that a low IRR is inevitable, and economic feasibility of a business is unlikely. However, when the target market is Japan where the prevalent electricity prices are much greater, that an IRR exceeding 65% is readily attainable. This is extremely attractive to potential investors, making capitalization of a venture likely.

The paper shows that the capital investment required for the system can be less than $1 per installed watt, contributing less than 0.02 $/KW-hr to the cost of energy provision. Since selling prices in feasible regions range from 0.18 to over 030 $/kW-hr, these costs are but a small fraction of the operating expenses. Thus a very large IRR is possible for such a business.     

The critical satellite technical issues of future pervasive broadband low-cost communication networks

Using a baseline design for a 30/20 GHz system, this paper addresses the critical technical issues of signal waveform design, projected spacecraft technology, satellite launch options and satellite cost. With DPCM (differential pulse-coded modulation) video signal encoding, 32 Mb/s user-to-user data rate per channel, 10% overhead, two orthogonal polarizations, and crosstalk loss limited to 1 dB, TFM (tamed frequency modulation—a type of QPSK) permits about 75 channels/GHz of frequency allocation, i.e. about twice the capacity possible with MSK or SFSK (types of QPSK). The BOM (beginning of mission) weight and power of a baseline 400-channel multi-beam satellite is about 1800 kg and 5000 W. Each 35 Mb/s channel can support 1–10 video channels depending on the signal processing at the ground terminals. These weight and power estimates assume hardened digital logic, composite material for a multibeam antenna structure, high-efficiency solar cells (45 W/kg), NiH₂ batteries and 10W/20 GHz power amplifiers. If more speculative late-1980s solar cell technology is assumed, then the BOM weight decreases about 10–15%. Using the Space Transportation System being developed by NASA, geosynchronous deployment options are shown as a function of communications capacity. Based on a cost model for large communication satellites, the total space segment cost of two active satellites and one spare would be about $485M. This system would have a peak capacity of 14 Gb/s or 400-35 Mb/s simultaneous one-way channels. Demonstration and/or initial system (ground plus space) costs would be significantly less. It is concluded that the above baseline satellite appears to be technically feasible in the late-1980s.     

An economically viable space power relay system

This paper describes and analyses the economics of a power relay system that takes advantage of recent technological advances to implement a system that is economically viable. A series of power relay systems are described which transport power ranging from 1250 to 5000 MW and distribute it to receiving sites at transcontinental distances. It is shown that, when offering electricity at prices competitive to those prevalent in developed cities in the USA, that a low IRR is inevitable, and economic feasibility of a business is unlikely. However, when the target market is Japan, where the prevalent electricity prices are much higher an IRR exceeding 65% is readily attainable. This is extremely attractive to potential investors, making capitalization of a venture likely. The paper shows that the capital investment required for the system can be less than $1 per installed watt, contributing less than 0.02 $ to the cost of energy provision. Since selling prices in feasible regions range from 0.18 to over 0.30 $, these costs are but a small fraction of the operating expenses. Thus a very large IRR is possible for such a business.     

Lunar oxygen: Economic benefit for transportation systems and in situ production by fluorination

The use of oxygen produced on the Moon—called “MOONLOX”—is considered as a propellant component for a reusable Earth-Moon transportation system consisting of an aeroassisted orbital transfer vehicle and a lunar bus for lunar descent/ascent. Conditions for economic benefit are discussed and the processing concept of a lunar oxygen plant based on fluorination is presented. It is shown that the necessary mass of supply from Earth for MOONLOX-production is an important parameter, which may not be neglected due to its strong influence on the economy. In the ideal case where no supplies from Earth are required a reduction of up to 50% in masses to be launched into low Earth orbit can be obtained for a typical lunar mission with use of MOONLOX compared to a reference scenario with Earth-derived propellant. Mass-saving decreases, however, significantly with increasing supply from Earth until a critical supply-rate is reached—measured in percentage of MOONLOX-mass produced and consumed—beyond which mass-saving and thus economically promising lunar oxygen production is no longer possible. This critical supply-rate depends on the scenario for MOONLOX-utilization and is much larger in the case of in situ use of MOONLOX on the lunar surface, e.g. as ascent propellant for the lunar bus, than in the case of export for complete refuelling of both space vehicles. The latter scenario therefore requires significantly more autonomy for MOONLOX-production. The reduction of masses to be transported into low Earth orbit and corresponding MOONLOX-consumption define for given specific Earth-to-LEO transportation costs an upper limit on MOONLOX-production costs beyond which economic benefit is not possible. Depending on the MOONLOX-utilization strategy this upper limit varies between 3000 and 55000 $/kg for current Earth-to-LEO transportation costs.     

Cost effectiveness of Spacelab experiments

Spacelab permits investigation in new seicntific disciplines like material processing, life sciences, chemistry, etc. The large mass and volume capabilities of Spacelab offer better possibilities for some areas of traditional space sciences like infrared astronomy, multi-spectral solar observations and large instruments for astronomical observations.Since free-flyers will require normally a new spacecraft development for each mission, the reusability of space qualified components and experiments will be a significant cost reduction factor over a long period. In the early phase of Spacelab utilisation, however, the scaling factor introduced by Spacelab utilisation, however, the scaling factor introduced by Spacelab results in higher payload development costs than originally appreciated.The costs of Spacelab utilisation are computed and compared with those of conventional free-flying satellites. The mission implementation costs and experiment development costs are shown for both cases. The Spacelab mission implementation costs are subdivided into NASA charges for the Standard Shuttle Mission, NASA charges to fly and operate Spacelab, the European costs of Spacelab payload integration and experiment development costs. In order to evaluate and compare mission implementation costs, the simple parameters are adopted of the cost per kg of experiments and the data collection-transmission capability of Shuttle/Spacelab and ESRO/ESA satellites. The mission implementation costs turn out to be very favourable for Spacelab. The experiment development costs, which are not included in the mission implementation costs, are compared for several free flyers with the corresponding development costs for several experiments of the first Spacelab payload. The comparison shows that the cost per kg of Spacelab experiment development is about five times less than of satellite experiments.     

Simulation model for the operation of space freighters

This paper makes the attempt to illustrate the need for a detailed operational analysis of future space transportation systems with the help of computer-based simulation models. The basic approach deemed suitable for such a systems simulation is explained in some detail. A reference program (100,000 Mg payload per year during 25 years) for a reference mission (heavy cargo transport to GEO for SPS construction) has been selected. A base-line launch vehicle (fully reusable ballistic all chemical three stage vehicle) has been defined, which is considered a serious applicant for such a mission. It was found that the take-off mass of this type of vehicle should be as large as practical from the viewpoint of cost-effectiveness. The example chosen has a GLOW of 10,000 Mg and lifts more than 100 Mg to GEO.With consideration of the operational parameters the simulation model evaluates the annual production rates, inventory of stages, utilization of facilities and operational cost, which amount within this frame of reference to about 96 $/kg net payload delivered to GEO in terms of 1980 dollars and contribute the main share to the total transportation cost.     

The future of human spaceflight.

After the Apollo Moon program, the international space station represents a further milestone of humankind in space, International follow-on programs like a manned return to the Moon and a first manned Mars Mission can be considered as the next logical step. More and more attention is also paid to the topic of future space tourism in Earth orbit, which is currently under investigation in the USA, Japan and Europe due to its multibillion dollar market potential and high acceptance in society. The wide variety of experience, gained within the space station program, should be used in order to achieve time and cost savings for future manned programs. Different strategies and roadmaps are investigated for space tourism and human missions to the Moon and Mars, based on a comprehensive systems analysis approach. By using DLR's software tool FAST (Fast Assessment of Space Technologies), different scenarios will be defined, optimised and finally evaluated with respect to mission architecture, required technologies, total costs and program duration. This includes trajectory analysis, spacecraft design on subsystem level, operations and life cycle cost analysis. For space tourism, an expected evolutionary roadmap will be described which is initiated by short suborbital tourism and ends with visionary designs like the Space Hotel Berlin and the Space Hotel Europe concept. Furthermore the potential space tourism market, its economic meaning as well as the expected range of the costs of a space ticket (e.g. $50,000 for a suborbital flight) will be analysed and quantified. For human missions to the Moon and Mars, an international 20 year program for the first decades of the next millennium is proposed, which requires about $2.5 Billion per year for a manned return to the Moon program and about $2.6 Billion per year for the first 3 manned Mars missions. This is about the annual budget, which is currently spend by the USA only for the operations of its Space Shuttle fleet which generally proofs the affordability of such ambitious programs after the build-up of the International Space Station, when corresponding budget might become again available.     

Radioisotope electric propulsion of sciencecraft to the outer solar system and near-interstellar space

Recent results are presented in the study of radioisotope electric propulsion as a near-term technology for sending small robotic sciencecraft to the outer Solar System and near- interstellar space. Radioisotope electric propulsion (REP) systems are low-thrust, ion propulsion units based on radioisotope electric generators and ion thrusters. Powerplant specific masses are expected to be in the range of 100 to 200 kg/kW of thrust power. Planetary rendezvous missions to Pluto, fast missions to the heliopause (100 AU) with the capability to decelerate an orbiter for an extended science program and prestellar missions to the first gravitational lens focus of the Sun (550 AU) are investigated.     

A practical approach to the disposal of highly toxic and long-lived spent nuclear fuel waste between Venus and Earth

Extraterrestrial disposal, while not the only alternative (there is at least one very safe terrestrial method), nevertheless assures definite and irreversible removal of the most toxic and long-lived waste from the biosphere. In the foreseeable future, there is far less need to retrieve nuclear waste for later use then to dispose of it in a publicly acceptable manner, beginning in the near future (middle 1980s). It is, therefore, more important to assure safety in the weakest link of this disposal process—transportation into near-Earth orbit—than to engage in a retrievable disposal beyond Earth. The disposal “site” should lie at minimum safe transfer energy level. Primary candidate is the space between Venus and Earth. The number of propulsion phases should be a minimum, preferably only one (out of near-Earth orbit). Lunar gravity assist can be helpful to achieve higher inclination of the heliocentric orbit relative to the ecliptic.Solidified spent fuel isotopes and actinides, sufficient to reduce the residual terrestrial waste to the radiation level of natural uranium deposits after 30–40 yr instead of 1000–1500 yr, is deposited into heliocentric orbits. Transportation systems, requirements, costs and the associated socio-economic benefit potentials of an environmentally more benign and a more vigorous nuclear power generation program are presented.Prior to solidification, an interim storage of 10 yr, following removal from the reactor, may be required. The Shuttle, with one Orbiter modified as Nuclear Waste Carrying Orbiter (NWCO) and an out of near-Earth orbit booster, provides a safe and economic transportation system at (1979$) disposal mission costs from surface to disposal orbit of less than $\text{0.5¢/}\text{kWhe or}\text{? 0.1¢/}\text{kWhe}$ (some 70% of overall exo-disposal cost), depending on level of orbital operations (this at basic Shuttle flight cost of $30M). The orbital operations mode can be selected according to lead time and costs, and can be advanced sequentially, lowering disposal costs while at the same time financing the development of orbital operations techniques needed also for other and larger-scale exo-industrial activities. An average of 10–15 disposal missions of the NWCO is required annually, at the given conditions, to service the spent fuel of 173 reactors 1978 in operation in W. Europe, Japan and North America.     

Data on and methodology for calculating space shuttle programme costs

This report presents data on the annual and total costs of NASA's Space Shuttle programme through fiscal year (FY) 1993. The total cost of the programme through FY 1993 is found to be $83.7 billion in 1992 dollars. This information has significance for pending policy decisions on the future of the Shuttle programme, its possible successors, and interrelated programmes, such as the Space Station.     

Recent advances in lunar base simulation

This article reports about the results of the latest computer runs of a lunar base simulation model. The lunar base consists of 20 facilities for lunar mining, processing and fabrication. The infrastructure includes solar and nuclear power plants, a central workshop, habitat and farm. Lunar products can be used for construction of solar power systems (SPS) or other spacecraft at several space locations. The simulation model evaluates the mass, energy and manpower flows between the elements of the system as well as system cost and cost of products on an annual basis for a given operational period. The 1983 standard model run over a fifty-years life cycle (beginning about the year 2000) was accomplished for a mean annual production volume of 78 180 Mg of hardware products for export resulting in average specific manufacturing cost of 8.4 $/kg and total annual cost of 1.25 billion dollars during the life cycle. The reference space transportation system uses LOX/LH₂ propulsion for which at the average 210 500 Mg LOX per year is produced on the moon. The sensitivity analysis indicates the importance of bootstrapping as well as the influence of market size, space transportation cost and specific resources demand on the mean lunar manufacturing cost. The option using lunar resources turns out to be quite attractive from the economical viewpoint. Systems analysis by this lunar base model and further trade-offs will be a useful tool to confirm this.     

European Sail Tower SPS concept

Based on a DLR-study in 1998/99 on behalf of ESA/ESTEC called “System Concepts, Architectures and Technologies for Space Exploration and Utilization (SE&U)” a new design for an Earth-orbiting Solar Power Satellite (SPS) has been developed. The design is called “European Sail Tower SPS” and consists mainly of deployable sail-like structures derived from the ongoing DLR/ESA solar sail technology development activity. Such a SPS satellite features an extremely light-weight and large tower-like orbital system and could supply Europe with significant amounts of electrical power generated by photovoltaic cells and subsequently transmitted to Earth via microwaves. In order to build up the sail tower, 60 units - each consisting of a pair of square-shaped sails - are moved from LEO to GEO with electric propulsion and successively assembled in GEO robotically on a central strut. Each single sail has dimensions of 150m × 150 m and is automatically deployed, using four diagonal light-weight carbon fiber (CFRP) booms which are initially rolled up on a central hub. The electric thrusters for the transport to GEO could also be used for orbit and attitude control of the assembled tower which has a total length of about 15 km and would be mainly gravity gradient stabilized. Employing thin film solar cell technology, each sail is used as a solar array and produces an electric power in orbit of about 3.7 MW_e. A microwave antenna with a diameter of 1 km transmits the power to a 10 km rectenna on the ground. The total mass of this 450 MW SPS is about 2100 tons. First estimates indicate that the costs for one kWh delivered in this way could compete with present day energy costs, if launch costs would decrease by two orders of magnitude. Furthermore, mass production and large numbers of installed SPS systems must be assumed in order to lower significantly the production costs and to reduce the influence of the expensive technology development. The paper presents the technical concept and an economic assessment as well as results of a recent solar sail deployment ground demonstration at DLR's facilities in Cologne.     

Small spacecraft “UNISAT” launched by the “start-1” launcher

The preliminary estimations show that the contemporary level of electronic and information engineering makes it possible to create a small s/c of 150–200 kg mass capable to solve both the problems of Earth remote sensing and many other applied and scientific problems orbiting the planets at 500–1000 km. In accordance with the fundamental criterion for choosing parameters of small multipurpose spacecraft the small UNISAT s/c has been created on the basis of a unified space platform. The design provides for s/c energetic, thermal and space-saving parameters satisfying the conditions for accommodation of various-purpose payload and a possibility of using relatively inexpensive and light launchers like “Start-1” mobile launch complexes. Space platform mass is 100–120 kg; permissible payloads (PL) mass is 40–80 kg; maximal average power consumption of the payload is up to 60 W; three-axes orientation accuracy up to 0.001 deg./s; s/c lifetime is not less than 3–5 years.     

一种气动式冲击响应谱试验系统的设计与分析

大多数航天器单机产品都要求进行冲击响应谱试验,试验量级普遍为1000g~2000g,甚至超过3000g。对于较大质量(如50 kg以上)产品高量级的冲击试验,振动台、传统的摆锤式或跌落式冲击试验台均很难满足相关要求。文章研究并设计了一套能够进行大质量受试产品高量级冲击响应谱试验的气动冲击试验系统。该系统利用压缩空气瞬间释放膨胀推动质量块加速撞击具有多阶固有频率的谐振板,通过谐振板被激起的响应模拟复杂的衰减正弦波。测试结果表明,系统空载时冲击谱量级达8000g,负载200 kg时可达5000g,时域曲线为振荡衰减波,持续时间小于10 ms。文章提出的气动式冲击响应谱试验系统设计方法可为此类冲击试验系统的设计提供参考和理论依据。     

Launch cost reduction by the reuse of the core stage of Ariane 5

The high cost of launching payloads into Earth orbit is a main limiting factor on the development of space. In order to reduce the high cost of launch, reuse of (parts of) the launch vehicle is needed. This study analyses the possibilities of recovering and reusing the core stage of Ariane 5. Recovery of the core stage sets demands on re-entry trajectory, attitude, stability, thermal protection, structural strength, terminal deceleration, salt water protection, recovery and refurbishment. All these subject areas require solutions to their individual problems. Added subsystems to the stage are defined and their mass is determined. These masses are used to determine the financial feasibility of the recovery concept, by weighing the payload demise and operational cost against the gains of reduced production cost. It is concluded that the recovery is technologically feasible, using a detachable ablative heat shield on the nose of the stage and a stabilisation device (an inflatable drag cone), a parachute system and an engine enclosure device. Total mass of these systems is 1320 kg, with financial savings amounting to $8.5 million per flight.     

The Hidden Costs of Reliability and Failure in Launch Systems

In comparing the costs of different launch vehicles, the possibility of the risk of failure is assumed to be accounted for by the cost of insurance. The satellite may be insured against loss during launch, and the launch services provider may offer a free relaunch. However, actual costs of reliability and failure extend beyond this. Each failure necessitates an investigation and a get well programme by the operating agency, while putting the operations team on hold until services can resume. A commercial operator may also lose customer revenue and actual customers through loss of confidence or unavailability. Such costs tend to be hidden, and not evaluated in assessing the effectiveness of a system, but count towards total costs. Failure investigations help to improve system reliability, but this could equally have been achieved by expenditure in development and qualification. Reusable launch vehicles will have different costs associated with reliability and failure. The relationship between reliability and cost, properly assessed, ought to influence the design of both expendable and reusable launch systems.     

分离模块异构航天器集群松散队形设计

针对近地轨道分离模块航天器集群,考虑航天器之间面质比差异,提出一种能在较长时间内维持两星之间相对距离的松散队形设计方法。该方法以相对偏心率向量和相对轨道倾角向量描述松散队形,以两颗卫星之间的轨道长半轴之差和相对相位角之差为优化变量,并且通过设置虚拟空间圆解决避碰问题,保证在考虑气动力摄动的情况下,异构航天器松散编队能够有较长时间满足距离要求。仿真研究表明,针对轨道高度为500 km的参考轨道,当两星面质比分别为0.003 m 2/kg和0.0032 m 2/kg时,两颗卫星之间的距离能够在50天内处于200 m到20 km之间。     

Success criteria for the next generation of space based multimedia systems

It has been almost three years since the great Geostationary Orbit Earth (GEO) Ka-band filing rush concluded at the FCC. The majority of filers have gone on to register their systems with the ITU, creating the most congested new Spectrum band at system startup that the world has ever seen. However, for all the hype, none of the GEO systems are thought yet to have begun to bend metal. All appear consumed with the requirement to find partners with whom to share the business risks, costs and rewards.

This Paper will examine the current state of progress of Kaband systems. Assessments will be made with respect to criteria such as: announced target markets and their appropriateness for the technology proposed; availability of desired technology to enable these systems to achieve their stated goals and a realistic assessment of when these technologies will be deployed; the availability of alternative terrestrial systems to address the same applications and their impact on the GEO Ka-band system target markets; and regulatory issues and paths to insure commercial access.     

氢仲-正转化释冷方案及冷量空间利用潜能分析

调研了氢仲-正转化释冷能力及研究现状,介绍了绝热、连续、等温三种转化方式的释冷潜能与工作特征。针对氢空间长期安全贮存,按照转化器布置位置与转化方式,提出四种冷量空间利用方案。研究发现：氢绝热转化在150 K时释冷量最大,为391 kJ/kg;等温转化在110 K时释冷量最大,为394 kJ/kg;连续转化在出口温度大于200 K后,释冷量稳定在491 kJ/kg。所提四种方案中,由于空间排气温度偏低,造成氢仲-正转化潜能无法充分释放,对贮箱绝热性能提升有限。相较于一次绝热转化,在蒸气冷却盘管内连续转化可较充分利用转化冷能,在氢的空间贮存应优先考虑。